The most widely used wing anti-icing system in commercial turbo-fan aircraft is the bleed air system. In this system, high pressure, high temperature air is fed from the compressor section of the engine and discharged through a perforated tube positioned within a leading edge chamber of the wing (or other aircraft structure) as a spray against the inner surface of the wing leading edge skin. This bleed air anti-icing system is generally an "on-off" type system without any modulating capability. The sizing criteria for the system is a maximum anti-icing effort during idle descent. Since bleed air temperature and pressure are very low at this power setting, the mass flow must be high. Therefore, the system is oversized and wasteful during less demanding conditions(thus resulting in a case of "overkill" under other conditions where icing can occur such as when the aircraft is climbing, with the engines operating at a much higher power setting).
The anti-icing air, after heating up the inside of the leading edge skin, is ducted overboard at a fairly high temperature. This dump loss can exceed 50% of the energy supplied.
Another consideration is that as turbo-fan engines are designed for higher propulsion efficiency, the gas generators become smaller and the penalties from bleed air become larger. On a modern turbo-prop engine, the gas generators are becoming so small that the amount of bleed air required for anti-icing and cabin pressurization/ventilation (Environmental Control System) can become prohibitive.
There are a number of other deicing and/or anti-icing systems in use other than the engine bleed air system described above. One such system is to utilize inflatable rubber boots at the aircraft leading edge portions, this being widely used for deicing on slower propeller airplanes. However, the boot does not provide a smooth enough surface for high performance aircraft. Also, the boot needs fairly frequent replacement due to erosion and aging.
Glycol anti-icing is another system, where the glycol is discharged over the surface areas where anti-icing is desired. However, this leaves a sticky residue on the wing which helps collect dust, so that frequent wing cleaning is required. Further, the cost and trouble of refueling glycol tanks are undesirable features.
Yet another approach is to utilize electric resistance heaters attached to the outside or inside of the leading edge skin. This can be a very energy efficient system. A drawback is that in the case of a defect, field repair can be difficult and there am be a lay up of an aircraft for an undesirably long period of time.
A search of the patent literature has disclosed a number of U.S. and foreign patents, these being the following.
U.S. Pat. No. 1,703,612 (Corousso) shows a variety of deicing configurations for a biplane. In FIGS. 10, 11, and 12, there is illustrated a distribution pipe 24 having an arcuate cross section, with exhaust nozzles 25 extending from the top and bottom edges of the pipe, so that hot air introduced into the pipe 24 escapes from these nozzles, and travels rearwardly along the wing. Thus, it would appear that the leading edge portion of the pipe 24 would be deiced (or have anti-icing) by reason of heat conduction through the forward portion of the pipe.
U.S. Pat. No. 2,318,233 (Keller) shows a variable pitch propeller for a piston powered aircraft. There is a double walled spinner at the front of the propeller shaft. The inner wall of the spinner has openings 9 through which hot air flows to exhaust over the exposed surface of the spinner at locations downstream of the further forward surface of the outer wall of the spinner.
U.S. Pat. No. 2,328,079 (Goodman) discloses a hollow boot which extends along and covers the leading edge of the wing. Hot air heats the boot by conduction, and this air then escapes through a slot at the rear of the boot to flow rearwardly over the wing surfaces.
U.S. Pat. No. 2,390,093 (Garrison) utilizes a porous sintered metal plate which forms the leading edge of the wing. Pressurized air carrying an atomized or vaporized deicing fluid moves through the porous metal plates to flow to the exterior surface of the leading edge and distribute the deicing fluid over the leading edge surface. The patent indicates that good results may be obtained with sintered metal sheets which pass air at the rate of 0.125 cubic feet per minute at a pressure drop of 39 inches of mercury. The patent also points out that experimental tests have shown that an air pressure of about 45 PSI is required to break up a heavy layer of ice when air alone is used, while a pressure of 10 PSI is sufficient for deicing when atomized or vaporized deicing fluid is mixed with the air. Thus, it would appear that the air is not used primarily to provide heat for deicing, but in one mode serves to carry the deicing fluid to the leading edge surface while in another mode the pressure of the air is utilized to break loose the ice which has already formed.
U.S. Pat. No. 2,482,720 (Sammons) discloses an anti-icing system for a turbo-engine. Hot air within a cavity heats an outer engine wall surface and an engine spinner by conduction. Then the slightly cooled air is exhausted outwardly through apertures in the inner engine inlet and the spinner to heat their respective surfaces at a further rearward location.
U.S. Pat. No. 2,625,010 (Clark) shows a gas turbine engine inlet anti-icing system which uses an auxiliary combustion chamber to create gasses for distribution to an engine inlet area during periods of low compressor bleed air availability. There are apertures in the interior of the engine inlet, used for exhausting the heating gas after this gas has heated the leading edges by conduction.
U.S. Pat. No. 2,630,965 (Greatrex et al) discloses a gas turbine engine inlet anti-icing system. The leading edges and exterior of the engine are heated by conduction through the skin from hot gasses circulated in an interior cavity. The slightly cooled gasses are then exhausted into the air stream from trailing edge apertures in stator or rotors.
U.S. Pat. No. 2,634,049 (Hodges) discloses a gas turbine engine inlet anti-icing system. The inlet guide vanes in FIG. 6 are shown having apertures or slots 56 which are formed on opposite sides of (and rearwardly of) the leading edge 55 of the guide vanes. The heated air is directed through the apertures 56 to flow over the upper and lower surfaces of the guide vanes, but not over the leading edges thereof.
U.S. Pat. No. 2,636,666 (Lombard) discloses a gas turbine engine where there is a forward wide mesh grid-like structure 17 extending across a forward portion of the engine inlet. Rearwardly of the grid 17, there is a second guard grid 19 of wire mesh which extends across the air intake guide to prevent any foreign matter, such as stones from entering the compressor and damaging the compressor blades. Hot gas is fed into the grid-like structure 17 which has outlets at the trailing edges of the members of the grid 17 so that the hot gas is delivered into the air intake duct 18 and thus reduce the possibility of ice formation on the guard grid 19 and other components.
U.S. Pat. No. 2,668,596 (Elliot) discloses a turbo-prop engine having the surface of the interior wall of the nacelle formed with apertures. Heated air is directed into the forward portion of the nacelle to heat the leading edge of the nacelle by conduction, after which the air is discharged through the interiorly directed apertures to flow rearwardly into the engine inlet.
U.S. Pat. No. 3,933,327 (Cook) provides for anti-icing of the inlet of a gas turbine engine. The hot gas heats the leading edge of the nacelle by conduction, with the slightly cooled gasses then being exhausted through a perforated acoustical member which is located rearwardly of the inlet lip.
U.S. Pat. No. 3,981,466 (Shah) discloses an anti-icing system for a gas turbine engine which provides heat by conduction from internal passages. The cooled anti-icing air is then utilized for heating in the interior space of the aircraft.
U.S. Pat. No. 4,099,691 (Swanson et al) discloses a boundary layer control system for aircraft where hot bleed air from the engine is directed to a number of manifolds located in a spanwise arrangement along the wing. A single row of air discharge openings (see FIGS. 9 and 10) are provided at an upper location adjacent to the leading edge of the wing to cause the boundary layer control air to flow upwardly along the upper wing surface aft of the leading edge. The patent is directed toward the thermal stress and aerodynamic problems associated with such boundary layer control systems.
U.S. Pat. No. 4,615,499 (Knowler)shows an improved "trombone" fitting in deploying leading edge slats.
British Patent Specification 504,360 (Spearpoint et al) discloses a lift increasing arrangement where high velocity air from the prop wash is collected and distributed spanwise along the wing. Then air is discharged upwardly and outwardly from a slot located slightly rearwardly of the leading edge so that the discharged air flows rearwardly along the upward surface of the wing. A heater in the duct provides anti-icing when required.
French Patent 972,392 (Greenly) discloses what appears to be an anti-icing method for the spinner and nacelle on a turbo-prop engine. A translation of this patent is not available, but it appears that the leading edge is heated by conduction, and then the hot gasses are exhausted from slots located at the interior of the nacelle rearwardly of the leading edge. It also appears that air is discharged from a single slot at the nose of the spinner toward a forward shield with this air then flowing so as to be radiating outwardly and rearwardly along the spinner.